Electric Flight Control System for Aircraft Elevators

ABSTRACT

The invention relates to an electric flight control system for aircraft elevators. According to the invention, the flight control system can be controlled in terms of load factor or rate of pitch. The inventive system comprises built-in protections in relation to load factor, incidence and pitch attitude.

The present invention relates to an electric flight control system forthe control of the elevators of an aircraft.

It is known that, in airplanes with mechanical controls, the stickdirectly controls an angle of deflection of the elevators, the amplitudeof this angle being proportional to the swing of said stick. Thus, byacting on said stick, the pilot exerts a piloting action whichmomentarily positions the airplane about its center of gravity or whichcontrols the attitude of said airplane. Such a piloting action isinstinctive for the pilot.

It is also known that the electric flight controls, with which numerousmodern airplanes are now equipped, make it possible to control anairplane by objective, that is to say to directly control a parameter,such as the load factor, by way of said stick, everything occurring asif the latter were graduated in terms of load factor. Such control isadvantageous since, by graduating the stick in terms of load factor andby limiting the extent of controllable load factor, one is sure topreserve the integrity of the airplane in any maneuver whatsoever.

On the other hand, control in terms of load factor is not veryinstinctive for the pilot, since control in terms of load factor makesit possible to guide the airplane in the long term rather than to pilotit at an instant. Specifically, commanding a zero load factor amounts toprescribing a constant aerodynamic slope. By virtue of the stick, it isthus possible to guide the airplane by controlling its trajectory,thereby making it possible to easily use said electric flight controlseither with a human pilot, or with an automatic pilot.

However, it results from the foregoing that, in very dynamic flightphases, for example in proximity to the ground, during which the controltasks are very short term, one is compelled to substitute with, forexample said control in terms of load factor control in terms of rate ofpitch.

Additionally, the known electric controls in terms of load factor cannotintegrate protection of the airplane in terms of longitudinal attitude.It is therefore necessary to append thereto a supplementaryattitude-based protection law and to interface the latter accuratelywith the electric flight control in terms of load factor, so as to avoidproblems in proximity to the ground, such as touchdown of the tail (ortail-strike).

Moreover, it is also impossible to protect the airplane in terms ofincidence solely through electric controls in terms of load factor. Itis therefore necessary, here again, to add a supplementaryincidence-based protection law and to undertake deft interfacing toavoid the risks of stalling of the airplane during standard maneuvers.

The object of the present invention is to remedy these drawbacks and itrelates to electric flight controls making it possible to control theaircraft in terms of load factor in certain flight conditions and interms of rate of pitch in others and integrating protections in terms ofload factor, incidence and longitudinal attitude.

For this purpose, according to the invention, the electric flightcontrol system for the control in terms of load factor of the elevatorsof an aircraft, said elevators being controlled by control meanscompelling said elevators to take a deflection position dependent on anelectrical signal δ_(mc) representative of the controlled value of theangle of deflection δ_(m) of said elevators, is noteworthy in that itcomprises:

-   -   first means of calculation for calculating, on the basis of an        electrical signal nz_(c) representative of a controlled value of        said load factor, a first electrical signal {dot over (γ)}_(c)        representative of the controlled value of the derivative, with        respect to time, of the aerodynamic slope γ of said aircraft;    -   a first constituent device, which:        -   is able to receive at its input said first electrical signal            {dot over (γ)}_(c);        -   comprises first means of protection able to maintain said            first electrical signal {dot over (γ)}_(c) between a minimum            value and a maximum value;        -   on the basis of said first electrical signal {dot over            (γ)}_(c) determines at least a second electrical signal            α_(c), representative of the corresponding controlled value            of the incidence α, and a third electrical signal θ_(c),            representative of the corresponding controlled value of the            longitudinal attitude θ;        -   comprises second means of protection able to maintain said            second electrical signal α_(c) between a minimum value and a            maximum value; and        -   delivers at least said third electrical signal θ_(c) to its            output; and    -   a second constituent device, which        -   is able to receive at its input at least said third            electrical signal θ_(c) or a fourth electrical signal θ_(d)            similar to said third electrical signal θ_(c);        -   comprises third means of protection able to maintain said            third or fourth electrical signal between a minimum value            and a maximum value; and        -   is able to deliver at its output a fifth electrical signal            which constitutes said signal δ_(mc), representative of the            corresponding controlled value of the angle of deflection            δ_(m) of said elevators.

Advantageously, the system in accordance with the invention comprisesfirst means of switching that can take:

-   -   either a first position for which the output of said first        constituent device is connected to the input of said second        constituent device, so that said third electrical signal θ_(c)        is then transmitted to said second constituent device;        -   or a second position for which the input of said second            device receives said fourth electrical signal θ_(d), similar            to said third electrical signal θ_(c) and produced on the            basis of a sixth electrical signal q_(d), representative of            a desired value for the rate of pitch q.

It is noted that, in a known manner, the load factor nz is equal toV.{dot over (γ)}/g in which expression V is the speed of the aircraft, gthe acceleration due to gravity and {dot over (γ)} the derivative of theaerodynamic slope. It is thus easy to transform the controlled loadfactor signal nz_(c) into said first electrical signal {dot over(γ)}_(c), since then {dot over (γ)}_(c) is equal to nz_(c).g/V.

Preferably, said first constituent device determines, in addition tosaid second electrical signal α_(c) and said third electrical signalθ_(c), a seventh electrical signal q_(c), representative of thecorresponding controlled value of the rate of pitch q, and said firstmeans of switching are able to transmit said seventh electrical signalq_(c) to said second constituent device.

By simplification, said first constituent device delivers, for theseventh electrical signal q_(c), an approximate value equal to that ofsaid first electrical signal {dot over (γ)}_(c).

In an advantageous embodiment of the electric flight control system inaccordance with the present invention, said system comprises:

-   -   an automatic pilot able to deliver a controlled load factor        signal nz_(c);    -   a manual piloting member able to deliver, by switching, either a        controlled load factor signal nz_(c), or said sixth electrical        signal q_(d), representative of a desired value for the rate of        pitch q; and    -   second means of switching for:        -   transmitting to said first constituent device either the            controlled load factor signal delivered by said automatic            pilot, or the controlled load factor signal delivered by            said manual piloting member;        -   or else transmitting said sixth electrical signal q_(d) to            first means of integration able to form the fourth            electrical signal θ_(d), representative of a desired value            for the attitude θ,    -    said first means of switching being able to transmit to said        second constituent device said fourth and sixth electrical        signals θ_(d) and q_(d), instead of said third and seventh        electrical signals θ_(c) and q_(c) produced by said first        constituent device.

Preferably, to determine said second electrical signal α_(c) on thebasis of the first signal {dot over (γ)}_(c), said first constituentdevice comprises second means of calculation calculating the expressionα_(c)=({dot over (γ)}_(c) −F _(γ))/G _(γ)in which F_(γ) and G_(γ) are functions of the state of the aircraft with$F_{\gamma} = {\frac{g \cdot {\cos(\gamma)}}{V} + {\frac{1}{2}{\frac{\rho}{m} \cdot V \cdot S \cdot {Cz}_{a = 0}}}}$and$G_{\gamma} = \left. {\frac{1}{2} \cdot \frac{\rho}{m} \cdot V \cdot S \cdot \frac{\partial{Cz}}{\partial\alpha}} \middle| {}_{\alpha = 0}{+ \frac{T}{m \cdot V}} \right.$where g is the acceleration due to gravity, γ the aerodynamic slope, Vthe speed of the aircraft, ρ the density of the air, m the mass of theaircraft, S the reference area of the aircraft, Cz_(α=0) the coefficientof lift of the aircraft for a zero incidence,$\frac{\partial{Cz}}{\partial\alpha}❘_{\alpha = 0}$the gradient of the aerodynamic coefficient of lift as a function of theincidence and T the thrust of the aircraft.

To form said third electrical signal θ_(c), said first constituentdevice may comprise second integrator means able to integrate saidseventh electrical signal q_(c) and a first summator for forming the sumof the integral delivered by said second integrator means and of saidsecond electrical signal α_(c).

Advantageously, said second constituent device, either on the basis ofsaid third electrical signal θ_(c) and of the seventh electrical signalq_(c) originating from said first constituent device, or on the basis ofsaid fourth signal θ_(d) and of said seventh signal q_(d) originatingfrom said manual piloting member, as well as current values q_(r) andθ_(r) of the rate of pitch q and of the longitudinal attitude θ,determines an eighth electrical signal {dot over (q)}_(c),representative of the corresponding controlled value of the pitchacceleration {dot over (q)}, then, on the basis of the eighth electricalsignal {dot over (q)}_(c), said second constituent device determinessaid fifth electrical signal δ_(mc).

Preferably, said second constituent device calculates said eighthelectrical signal {dot over (q)}_(c) through the relation{dot over (q)} _(c) =K1.θ_(v) −K2.θ_(r) +K3.q _(v) −K4.q _(r)where θ_(v) is said third or fourth electrical signal, θ_(r) the currentvalue of the longitudinal attitude θ, q_(v) said sixth or seventhelectrical signal, q_(r) the current value of the rate of pitch q, K1,K2, K3 and K4 being constant coefficients. Moreover, to determine saidfifth electrical signal δ_(mc) on the basis of said eighth electricalsignal {dot over (q)}_(c), said second constituent device advantageouslycomprises third means of calculation calculating the expressionδ_(mc)=({dot over (q)} _(c) −F _(q))/G _(q)in which F_(q) and G_(q) are functions of the state of the aircraft with$F_{q} = {{\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot {Cm}_{{\delta\quad m} = 0}} + {\frac{1}{I_{\gamma}} \cdot T \cdot b \cdot {\cos(\tau)}}}$and$G_{q} = \left. {\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot \frac{\partial{Cm}}{{\partial\delta}\quad m}} \right|_{{\delta\quad m} = 0}$where I_(y) is the pitch inertia, ρ the density of the air, V the speedof the aircraft, S the reference area of the aircraft, l the referencelength of the aircraft, Cm_(δm=0) the coefficient of pitch, T thethrust, b the lever arm of the engines, τ the angle of longitudinal trimof the engines and$\frac{\partial{Cm}}{{\partial\delta}\quad m}❘_{{\delta\quad m} = 0}$the effectiveness of the elevators.

The figures of the appended drawing will elucidate the manner in whichthe invention may be embodied. In these figures, identical referencesdesignate similar elements.

FIG. 1 is the schematic diagram of an exemplary embodiment of theelectric flight control system in accordance with the present invention.

FIG. 2 shows the diagram of an integration device used in the system ofFIG. 1,

FIG. 3 is the schematic diagram of the first constituent device of thesystem of FIG. 1.

FIG. 4 illustrates the gain curve of a device of FIG. 3.

FIG. 5 is the schematic diagram of the second constituent device of thesystem of FIG. 1.

FIGS. 6 and 7 are diagrams illustrating the nature of the means ofcalculation of the second constituent device calculating the eighthelectrical signal {dot over (q)}_(c).

In the diagrammatic exemplary embodiment of the electric flight controlsystem for aircraft, in accordance with the present invention and shownin FIG. 1, there is represented an automatic pilot 1, a stick system 2,directional control surfaces 3 and a device 4 for controlling thedeflection of said control surfaces.

In this example, it is assumed that the automatic pilot 1 was able toemit, on its output terminal 7, an electrical control signal nz_(c)corresponding to a controlled value of the load factor, while the sticksystem 2 was able to produce, on its output terminal 24, an electricalcontrol signal corresponding, by switching, either to a desired valueq_(d) of the rate of pitch, or to a controlled value nz_(c) of the loadfactor. Moreover, there is provided, on the one hand, means ofcalculation 5 able to calculate, on the basis of the controlled valuenz_(c), an electrical signal {dot over (γ)}_(c), representative of acontrolled value of the derivative, with respect to time, of theaerodynamic slope γ and, on the other hand, means of integration 6, ableto produce, on the basis of said desired value q_(d), an electricalsignal θ_(d), representative of a desired value of the longitudinalattitude θ.

The means of calculation 5 exploit the relationnz=V.{dot over (γ)}/glinking the load factor nz and the derivative {dot over (γ)} of theaerodynamic slope γ, V being the speed of the aircraft and g theacceleration due to gravity. Thus, the means of calculation 5, to whichthe speed V and the acceleration due to gravity g are addressed andwhich receive the controlled value nz_(c), determine the quantity {dotover (γ)}_(c) through the expression{dot over (γ)}_(c) =nz _(c) .g/V.

The integration means 6, for their part, comprise an input terminal 8and two output terminals 9 and 10. Between the terminals 8 and 9 isdisposed an integrator 11, while a link 12 connects the terminals 8 and10 (see FIG. 2). Thus, when the desired value q_(d) of the rate of pitchis applied to the input terminal 8, there are obtained, by integrationby the integrator 11, the desired value of longitudinal attitude θ_(d)on the output terminal 9 and the desired value of rate of pitch q_(d) onthe terminal 10.

This being previously described, it is seen in FIG. 1 that the exemplaryelectric flight control system according to the invention, which isrepresented therein, comprises:

-   -   a first constituent device 13, comprising an input terminal 14,        upstream of which are disposed the means of calculation 5, and        two output terminals 15 and 16;    -   a second constituent device 17, comprising two input terminals        18 and 19 and an output terminal 20, the latter output terminal        controlling the device 4;    -   a switch 21, able to connect the input terminal 27 of the means        of calculation 5, either to the output terminal 7 of the        automatic pilot 1, or to a terminal 22;    -   a switch, 23 able to connect the output terminal 24 of the stick        system 2, either to said terminal 22, or to the input terminal 8        of the integration means 6, said switch 23 controlling moreover        the switching between the pitch rate signal q_(d) and the signal        nz_(c) at said output terminal 24; when the switch 23 connects        the terminals 24 and 8, the output signal of the stick system 2        is the signal q_(d); conversely, when the switch 23 connects the        terminals 24 and 22, the output signal of the stick system 2 is        the signal nz_(c);    -   a switch 25, able to connect the input terminal 18 of the second        device 17, either to the output terminal 15 of the first device        13, or to the output terminal 9 of the integration means 6; and    -   a switch 26, able to connect the input terminal 19 of the second        device 17, either to the output terminal 16 of the first device        13, or to the output terminal 10 of the integration means 6.

From the foregoing, it may therefore been seen that:

-   -   when the switches 21, 25 and 26 connect respectively the        terminals 27 and 7, 18 and 15 and 19 and 16, the elevators 3 are        controlled on the basis of the automatic pilot 1, by way of the        first and second devices 13 and 17, the control parameter being        the controlled value nz_(c) of the load factor;    -   when the switches 21, 23, 25 and 26 connect respectively the        terminals 24 and 22, 22 and 27, 18 and 15 and 19 and 16, the        elevators 3 are controlled on the basis of the stick system 2,        by way of the first and second devices 13 and 17, the control        parameter being the controlled value nzC of said load factor;        and    -   when the switches 23, 25 and 26 connect respectively the        terminals 24 and 8, 18 and 9 and 19 and 10, the elevators 3 are        controlled on the basis of the stick system 2, by way of the        second device 17 alone, the control parameter being the desired        rate of pitch q_(d).

Represented in FIG. 3 is an exemplary embodiment for the firstconstituent device 13. In this exemplary embodiment, the first device 13comprises:

-   -   a protection device 30, for example of the voter type, receiving        the signal {dot over (γ)}_(c) of the means of calculation 5 on        the terminal 14 unable to protect this signal between a maximum        value {dot over (γ)}_(max) and a minimum value {dot over        (γ)}_(min);    -   a calculator 31 receiving the signal {dot over (γ)}_(c) of the        protection device 30, as well as a plurality of cues 32 and        calculating (in the manner described hereinafter) a controlled        incidence signal α_(c);    -   a protection device 33, for example of the voter type, receiving        the controlled incidence signal α_(c) and able to protect this        signal between a maximum value α_(max) and a minimum value        α_(min);    -   a link 34 between the output of the protection device 30 and the        output terminal 16, said link 34 comprising a device 35 of gain        K(α) varying as a function of the incidence α; it is noted that        at the output of the protection device 30 there appears the        signal {dot over (γ)}_(c), which is equal to the difference        between the controlled rate of pitch q_(c) and the derivative of        the controlled incidence {dot over (α)}_(c) (in fact, θ=α+γ).        Now, the derivative of the controlled incidence {dot over        (α)}_(c) is generally very noisy, so that it is preferable to        neglect it; hence the derivative of the controlled slope {dot        over (γ)}_(c) is used to estimate an approximate value of the        controlled rate of pitch q_(c). As shown in FIG. 4, the gain        K(α) is always equal to 1, except in the neighborhood of α_(min)        and of α_(max), where it decreases to zero. Protection in terms        of incidence is thus obtained;    -   a device 36 of gain 1−K(α), receiving the current aerodynamic        slope cue γ_(r);    -   a summator 37 adding together the output signals of the devices        35 and 36 and addressing the resultant signal thereof to the        terminal 16; this resultant signal is therefore {dot over        (γ)}_(c), when α is far from α_(min) and from α_(max) and γ_(r)        when α is equal to α_(min) or to α_(max);    -   an integrator 38, integrating the signal in the link 34 and        therefore forming a controlled slope γ_(c);    -   a summator 39, adding together said controlled slope γ_(c) given        by the integrator 38 and the controlled incidence α_(c),        originating from the protection device 33 to obtain the        controlled attitude θ_(c) and address it to the terminal 15,        possibly by way of a supplementary incidence-based protection        device 40, involving the current values α_(r) and θ_(r) of the        incidence and of the attitude.

The manner of operation of the calculator 31 is based on the equationfor the lift which may be written{dot over (γ)}=F _(γ) +G _(γ).αin which expression F_(γ) and G_(γ) are functions of the state of theaircraft with$F_{\gamma} = {\frac{g \cdot {\cos(\gamma)}}{V} + {\frac{1}{2}{\frac{\rho}{m} \cdot V \cdot S \cdot {Cz}_{a = 0}}}}$and$G_{\gamma} = \left. {\frac{1}{2} \cdot \frac{\rho}{m} \cdot V \cdot S \cdot \frac{\partial{Cz}}{\partial\alpha}} \middle| {}_{\alpha = 0}{+ \frac{T}{m \cdot V}} \right.$where g is the acceleration due to gravity, γ the aerodynamic slope, Vthe speed of the aircraft, ρ the density of the air, m the mass of theaircraft, S the reference area of the aircraft, Cz_(α=0) the coefficientof lift of the aircraft for a zero incidence,$\frac{\partial{Cz}}{\partial\alpha}❘_{\alpha = 0}$the gradient of the aerodynamic coefficient of lift as a function of theincidence and T the thrust of the aircraft.

The cues 32 received by the calculator 31 therefore consist of theparameters g, γ, V, ρ, m, S, Cz_(α=0),${\frac{{\partial C}\quad z}{\partial\alpha}❘_{\alpha = 0}}\quad$and T, available on board the aircraft and making it possible tocalculate F_(γ) and G_(γ), after which the calculator 31 calculatesα_(c) by the expressionα_(c)=({dot over (γ)}_(c) −F _(γ))/G _(γ).

Thus, on the terminals 15 and 16 of the first constituent device 13there appear respectively the controlled attitude θ_(c) and thecontrolled rate of pitch q_(c).

As was mentioned above, the controlled attitude signal θ_(v) appearingon the input terminal 18 of the second constituent device 17 is formedeither by the signal θ_(c), originating from the output terminal 15 ofthe first constituent device 13, or by the signal θ_(d), originatingfrom the means of integration 6.

Likewise, the controlled rate of pitch signal q, appearing on the inputterminal 19 of said second constituent device 17 is formed either by thesignal q_(c), originating from the output terminal 16 of the firstconstituent device 13, or by the signal q_(d) originating from the meansof integration 6.

In the second constituent device 17, the controlled longitudinalattitude signal θ_(v) is addressed to a protection device 41, forexample of the voter type, able to maintain this signal between aminimum value θ_(min) and a maximum value θ_(max). The controlled rateof pitch signal q_(v), for its part, is addressed to a gain device 42,whose gain K(θ) is always equal to 1, except when θ is in theneighborhood of θ_(min) or of θ_(max), for which values K(θ) is equal tozero (the curve of the gain K(θ) as a function of θ is similar to thatshown in FIG. 4 in regard to the variation of K(α) as a function of α).

On the basis of the values θ_(v) (thus protected) and q_(v) (passed intothe device 42), the constituent device 17 comprises means forcalculating the derivative with respect to time of q_(v) representativeof the controlled value {dot over (q)}_(c) of the pitch acceleration.These means of calculation comprise:

-   -   a gain device 43, of gain K1, receiving the signal θ_(v) of the        protection device 41;    -   a gain device 44, of gain K2, receiving the current value θ_(r)        of the longitudinal attitude;    -   a gain device 45, of gain K3, receiving the signal q_(v) of the        device 42;    -   a gain device 46, of gain K4, receiving the current value q_(r)        of the rate of pitch q;    -   a subtractor 47 for calculating the difference        K1.θ_(v)−K2.θ_(r);    -   a subtractor 48 for calculating the difference        K3.q_(v)−K4.q_(r); and    -   an adder 49 for computing the sum        {dot over (q)} _(c) =K1.θ_(v) −K2.θ_(r) +K3.q _(v) −K4.q _(r).

In it noted that, by Laplace transformation, this sum may be writtens ².θ_(r) =K1.θ_(v) −K2.θ_(r) +K3.sθ _(v) −K4.s.θ _(r)in which expression s is the Laplace operator, so thatθ_(r)/θ_(v)=(τ.s+1).ω²/(s ²+2.z.ω.s+ω ²)taking K1=K2=ω², K3=τ/ω² and K4−2.z.ω.

The elements 43 to 49 therefore behave like a second order filter, ofnatural angular frequency ω and damping z, with a first order phaseadvance with time constant equal to τ.

Represented respectively in FIGS. 6 and 7 are the indicial response andthe pursuit response of such a filter, as a function of time t.

The second constituent device 17 comprises moreover a calculator 50receiving the controlled value {dot over (q)}_(c) of the pitchacceleration, formulated by the elements 43 to 49, as well as aplurality of cues 51, and calculating the signal δ_(mc).

The manner of operation of the calculator 50 is based on the fact that,in a known fashion, the acceleration of the pitch {dot over (q)} of anaircraft is an affine function of the angle of deflection δ_(m) of theelevators 3, which may be written{dot over (q)}= F _(q) +G _(q).δ_(m)in which expression$F_{q} = {{\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot {Cm}_{{\delta\quad m} = 0}} + {\frac{1}{I_{\gamma}} \cdot T \cdot b \cdot {\cos(\tau)}}}$and$G_{q} = \left. {\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot \frac{\partial{Cm}}{{\partial\delta}\quad m}} \right|_{{\delta\quad m} = 0}$where I_(y) is the pitch inertia, ρ the density of the air, V the speedof the aircraft, S the reference area of the aircraft, l the referencelength of the aircraft, Cm_(δm=0) the coefficient of pitch, T thethrust, b the lever arm of the engines, τ the angle of longitudinal trimof the engines and$\frac{\partial{Cm}}{{\partial\delta}\quad m}❘_{{\delta\quad m} = 0}$the effectiveness of the elevators.

Thus, the cues 51 comprise I_(γ), ρ, V, S, l, Cm_(δm=0), T, b, τ and$\frac{\partial{Cm}}{{\partial\delta}\quad m}❘_{{\delta\quad m} = 0}$and the calculator 50 calculates the expressionδ_(mc)=({dot over (q)} _(c) −F _(q))/G _(q).

It is thus seen that, by virtue of the present invention, one obtains asystem of longitudinal electric flight controls with integratedprotections in terms of load factor, incidence and longitudinalattitude, the piloting parameter being able to be, by choice, the loadfactor or the rate of pitch.

1-10. (canceled)
 11. An electric flight control system for the controlin terms of load factor of the elevators (3) of an aircraft, saidelevators being controlled by control means (4) compelling saidelevators to take a deflection position dependent on an electricalsignal δ_(mc) representative of the controlled value of the angle ofdeflection δ_(m) of said elevators (4), wherein it comprises first meansof calculation for calculating, on the basis of an electrical signal nzcrepresentative of a controlled value of said load factor, a firstelectrical signal {dot over (γ)}_(c) representative of the controlledvalue of the derivative, with respect to time, of the aerodynamic slopeγ of said aircraft; a first constituent device (13), which: is able toreceive at its input (14) said first electrical signal {dot over(γ)}_(c); comprises first means of protection (30) able to maintain saidfirst electrical signal {dot over (γ)}_(c) between a minimum value and amaximum value; on the basis of said first electrical signal {dot over(γ)}_(c), determines at least a second electrical signal α_(c),representative of the corresponding controlled value of the incidence α,and a third electrical signal θ_(c), representative of the correspondingcontrolled value of the longitudinal attitude θ; comprises second meansof protection (33) able to maintain said second electrical signal α_(c)between a minimum value and a maximum value; and delivers at least saidthird electrical signal θ_(c) to its output (15); and a secondconstituent device (17), which is able to receive at its input (18) atleast said third electrical signal θ_(c) or a fourth electrical signalθ_(d) similar to said third electrical signal θ_(c); comprises thirdmeans of protection able to maintain said third or fourth electricalsignal between a minimum value and a maximum value; and is able todeliver at its output (20) a fifth electrical signal which constitutessaid signal δ_(mc), representative of the corresponding controlled valueof the angle of deflection δ_(m) of said elevators (4).
 12. The systemas claimed in claim 11, wherein it comprises first means of switching(25, 26) that can take: either a first position for which the output ofsaid first constituent device is connected to the input of said secondconstituent device, so that said third electrical signal θ_(c) is thentransmitted to said second constituent device; or a second position forwhich the input of said second device receives said fourth electricalsignal θ_(d), similar to said third electrical signal θ_(c) and producedon the basis of a sixth electrical signal q_(d), representative of adesired value for the rate of pitch q.
 13. The system as claimed inclaim 11, wherein said first constituent device (13) determines, inaddition to said second electrical signal α_(c) and said thirdelectrical signal θ_(c), a seventh electrical signal q_(c),representative of the corresponding controlled value of the rate ofpitch q, and said first means of switching are able to transmit saidseventh electrical signal q_(c) to said second constituent device (17).14. The system as claimed in claim 13, wherein said first constituentdevice (13) delivers, for the seventh electrical signal q_(c), aapproximate value equal to that of said first electrical signal {dotover (γ)}_(c).
 15. The system as claimed in claim 12, wherein itcomprises: an automatic pilot (1) able to deliver a controlled loadfactor signal nz_(c); a manual piloting member (2) able to deliver, byswitching, either a controlled load factor signal nz_(c) or said sixthelectrical signal q_(d), representative of a desired value for the rateof pitch q; and second means of switching (21, 23) for: transmitting tosaid first constituent device (13) either the controlled load factorsignal delivered by said automatic pilot, or the controlled load factorsignal delivered by said manual piloting member; or else transmittingsaid sixth electrical signal q_(d) to first means of integration (6)able to form the fourth electrical signal θ_(d), representative of adesired value for the attitude θ, said first means of switching (25, 26)being able to transmit to said second constituent device (17) saidfourth and sixth electrical signals θ_(d) and θ_(d), instead of saidthird and seventh electrical signals θ_(c) and q_(c) produced by saidfirst constituent device.
 16. The system as claimed in claim 11,wherein, to determine said second electrical signal α_(c) on the basisof the first signal {dot over (γ)}_(c), said first constituent devicecomprises second means of calculation (31) calculating the expressiona _(c)=({dot over (γ)}_(c) −F _(γ))/G _(γ) in which F_(γ) and G_(γ) arefunctions of the state of the aircraft with$F_{\gamma} = {\frac{g \cdot {\cos(\gamma)}}{V} + {\frac{1}{2}{\frac{\rho}{m} \cdot V \cdot S \cdot {Cz}_{a = 0}}}}$and$G_{\gamma} = \left. {\frac{1}{2} \cdot \frac{\rho}{m} \cdot V \cdot S \cdot \frac{\partial{Cz}}{\partial\alpha}} \middle| {}_{\alpha = 0}{+ \frac{T}{m \cdot V}} \right.$where g is the acceleration due to gravity, γ the aerodynamic slope, Vthe speed of the aircraft, ρ the density of the air, m the mass of theaircraft, S the reference area of the aircraft, Cz_(α=0) the coefficientof lift of the aircraft for a zero incidence,$\frac{\partial{Cz}}{\partial\alpha}❘_{\alpha = 0}$ the gradient of theaerodynamic coefficient of lift as a function of the incidence and T thethrust of the aircraft.
 17. The system as claimed in claim 13, whereinsaid first constituent device comprises second integrator means (38)able to integrate said seventh electrical signal q_(c) and a firstsummator (39) for forming the sum of the integral delivered by saidsecond integrator means (38) and of said second electrical signal αc, soas to form said third electrical signal θc.
 18. The system as claimed inclaim 15, wherein said second constituent device (17), either on thebasis of said third electrical signal θ_(c) and of the seventhelectrical signal q_(c) originating from said first constituent device(13), or on the basis of said fourth signal θ_(d) and of said sixthsignal q_(d) originating from said manual piloting member, as well ascurrent values q_(r) and θ_(r) of the rate of pitch q and of thelongitudinal attitude θ, determines an eighth electrical signal {dotover (q)}_(c), representative of the corresponding controlled value ofthe pitch acceleration {dot over (q)}, then, on the basis of this eighthelectrical signal {dot over (q)}_(c) said second constituent devicedetermines said fifth electrical signal δ_(mc).
 19. The system asclaimed in claim 18, wherein said second constituent device calculatessaid eighth electrical signal {dot over (q)}_(c), through the relation{dot over (q)} _(c) =K1.θ_(v) −K2.θ_(r) +K3.q _(v) −K4.q _(r) whereθ_(v) is said third or fourth electrical signal, θ_(r) the current valueof the longitudinal attitude θ, q_(v) said sixth or seventh electricalsignal, q_(r) the current value of the rate of pitch q, K1, K2, K3 andK4 being constant coefficients.
 20. The system as claimed in claim 18,wherein, to determine said fifth electrical signal δ_(mc) on the basisof said eighth electrical signal {dot over (q)}_(c), said secondconstituent device comprises third means of calculation calculating theexpressionδ_(mc)=({dot over (q)} _(c) ,−F _(q))/G _(q) in which F_(q) and G_(q)are functions of the state of the aircraft with$F_{q} = {{\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot {Cm}_{{\delta\quad m} = 0}} + {\frac{1}{I_{\gamma}} \cdot T \cdot b \cdot {\cos(\tau)}}}$and$G_{q} = \left. {\frac{1}{I_{y}} \cdot \frac{1}{2} \cdot \rho \cdot v^{2} \cdot S \cdot \ell \cdot \frac{\partial{Cm}}{{\partial\delta}\quad m}} \right|_{{\delta\quad m} = 0}$where l_(γ) is the pitch inertia, ρ the density of the air, V the speedof the aircraft, S the reference area of the aircraft, l the referencelength of the aircraft, Cm_(δm=0) the coefficient of pitch, T thethrust, b the lever arm of the engines, τ the angle of longitudinal trimof the engines and$\frac{\partial{Cm}}{{\partial\delta}\quad m}❘_{{\delta\quad m} = 0}$the effectiveness of the elevators